JPH02102894A - Heat protective system - Google Patents

Abstract

PURPOSE: To provide a resistant, lightweight and reusable thermal protection and structure system suitable for a hypersonic aerospace vehicle by integrally connecting an outer sheet formed of a ceramic matrix to a rigid insulator core containing porous foamed ceramic, and connecting the core to an inner support structure. CONSTITUTION: A thermal protection system is constituted of an outer face sheet or an external shell 10 for forming a hard resistant outer front surface, a rigid insulator core 12 to which the cuter face sheet 10 is integrally connected, and a main load carrier or a structural member 15 to be attached to the insulator core 12 by bonding. The structural member 15 is provided with an inner face sheet 16 and connected to the other structural element 14. The outer face sheet 10 is a ceramic matrix basically formed of high-temperature ceramic, and preferred ceramic preferable to be used is silicon carbide. The rigid insulator core 12 is formed of foamed ceramic, and a specified ceramic material to be used is silicon carbide of silicon nitride.

Description

BACKGROUND OF THE INVENTION This invention relates to systems for providing thermal resistance or protection to hypersonic aerospace vehicles, and more particularly to systems for providing thermal resistance or protection to hypersonic aerospace vehicles, and more particularly to systems for providing thermal resistance or protection to such vehicles.
s/TPS).

In the design of thermal protection systems for aerospace vehicles, such systems should not transfer excessive heat to the basic vehicle structure, have low ff1l, and produce low thermal stresses. Current thermal protection system concepts use large numbers of ceramic tiles as reusable surface insulation,
Numerous joints and structures are subject to differential thermal expansion which results in considerable undesirable weight. Such tiles have the further disadvantage of being weak, brittle, susceptible to surface cracking, and labor intensive both for production and maintenance.

Various thermal protection systems have been developed in the prior art to overcome these problems. Accordingly, US Patent No. 4.344.591 is directed to a multi-wall thermal protection system that replaces prior art tile systems. This patent uses a hard exterior surface panel concept. In one embodiment, a multi-wall panel is formed by alternating layers of recessed and flat titanium alloy foil sheets and beaded scarfed edge seals. Additional embodiments use an intermediate fibrous insulation in the sandwich panels, and a third embodiment uses a frog coated columbium waffle as the outer panel skin and an intermediate protection of a fibrous layer.

The panels of this patent are attached to load-bearing channels that are part of the fuselage by deflecting clips to compensate for thermal loads.

Other examples of thermal insulation and protection systems are found in Patent Nos. 3.177.811; 3,793°861;
．． 955.034; No. 4.173゜187: No. 3.23
6.476; TS3,920°339; No. 4.112,
179.

It is therefore an object of the present invention to provide a new and improved thermal protection system for hypersonic vehicles.

Another object of this invention is to provide a durable, lightweight, reusable thermal protection and structural system for high speed aerospace vehicles.

A further object of this invention is to provide improved cost for high speed aerospace vehicles that are easily constructed and maintained and are integrally formed with airframe components such as the aerospacecraft's fuselage, wings or vertical tail. The objective is to provide an effective thermal protection or insulation system.

SUMMARY OF THE INVENTION The above objects and advantages are achieved in accordance with the present invention by providing an integral structure and thermal protection system (Is/TPS) for a hypersonic aircraft comprising essentially a hard resistant exterior surface member, a rigid insulator core. The external surface member is integrally connected to the insulator core and the internal primary load carrier structure is suitably connected or attached to the insulator core.

According to one embodiment, the exterior surface member is a silicon carbide material, the rigid insulator core is foamed silicon carbide, and the primary load carrier structure is comprised of a graphite epoxy composite.

This arrangement functions as a crimp, with the thermal protection system supporting and stabilizing the primary load carrier structure. The insulator core of the thermal protection system is connected to the load carrier structure or components thereof by a suitable compatible process, such as bonding, brazing, or chemical vapor deposition.

The hard exterior surface of the thermal protection system of this invention resists handling and damage from foreign objects. The system of the invention resists oxidation even at the highest temperatures encountered and is waterproof. The Is/TPS of this invention is lightweight due to the use of thin exterior surfaces and a low density insulator core that aids the primary structure in load propagation. The system of the present invention is reusable because continuous operation of the aerospace vehicle does not degrade or dissipate the thermal protection system.

DETAILED DESCRIPTION OF THE INVENTION AND PREFERRED EMBODIMENTS Referring to FIG. 1 of the drawings, there is illustrated a monolithic construction and thermal protection system according to the invention, the arrangement of which includes an outer face sheet or outer shell forming a hard, resistant outer surface. , a rigid insulator core 1 to which an outer facesheet 10 is integrally connected.
2, and a primary load carrier or structural member 15 suitably connected or attached to the insulator core 12, for example by bonding. Structural member 15 has an interior facesheet 16 that is integral with or coupled to another component 14 of structural member 15 .

The outer facesheet or outer shell 10 is a ceramic matrix consisting essentially of high temperature ceramic. The preferred ceramic used is silicon carbide.

An example of such a ceramic matrix is Nextel (
Nextel) in the form of fiber-reinforced silicon carbide. Nextel fiber is alumina/silica/boria
) It is a fiber. Chemical vapor deposition can be used to deposit additional silicon carbide into the fiber if desired.

The outer facesheet or shell 10 can also be in the form of a coating deposited on heat treated fibers, such as Nextel. Thus, a relatively smooth outer coating of silicon carbide can be deposited on such fibers by vapor deposition of silicon carbide, which becomes an integral part of the facesheet.

The thickness of the outer facesheet 10 can be on the order of 0.01' when formed essentially of a ceramic matrix and the vapor-deposited silicon carbide used to form the outer facesheet 10. The thickness of the coating can be on the order of about 0.001'. The outer facesheet 10 generally has a thickness of about 0.01' to about 0.01'.
06'. It is preferred that the outer facesheet 10 be as thin as possible consistent with aerodynamic, thermal and structural loads, as well as ease and safety of handling.

Rigid insulator core 12 is formed from foamed ceramic. Particular ceramic materials that can be used are silicon carbide or silicon nitride.

An insulator core 12 is built directly onto the outer facesheet 10. Generally, insulator core 12, for example in the form of foamed silicon carbide, is deposited on facesheet 10 by chemical vapor deposition of silicon carbide on the facesheet.

Insulator core 12 is thus integrally attached to outer facesheet 10.

Foam core 12 is porous. Therefore, it is desirable to have a core that is as light as possible for most applications. Ceramic core 12 has a density of approximately 0.0 mm per cubic foot. 5 to 10 bonds, preferably about 0.000000000000 bonds per cubic foot. It can range from 5 to 1°5 bonds. When deposited by chemical vapor deposition as described above, the ceramic insulator core 12 has a bubble-like profile.

The thickness of the core 12 is a function of the amount of insulation desired, the length of time the aerospace vehicle is expected to be at an elevated temperature, and the nature of the desired temperature differential between the exterior and interior of the vehicle. Therefore, the selected ceramic foam material, foam density, and its thickness are the criteria that determine the temperature difference between the outside or inside of the vehicle by the end of its mission flight.

Generally, the thickness of the insulator core 12 is about 0.5' to about 6'
can span a range of

The primary structural material, generally indicated at 15, can be comprised of materials including composites, metals, and metal matrices including fiber reinforcement. Material selection for major structures, such as airframe components, depends on geographically varying external and internal temperature tolerances.

For example, composite materials that can be used as the primary structural material of the airframe can include graphite epoxy and graphite polyimide. Instead of graphite fibers, the reinforcing fibers of such composites can include boron fibers, such as boron epoxy and boron polyimide composites, due to their high elastic modulus.

The metals used as main structural materials are aluminum and aluminum alloys, such as aluminum lithium,
and titanium and titanium alloys, such as titanium aluminide.

Metal matrix materials that can be used as primary structural materials include beryllium, aluminum, and aluminum alloys, such as aluminum lithium, titanium and its alloys, such as titanium aluminide, steel, molybdenum alloys, and nickel alloys such as Inconel. Such metals and metal alloys include graphite, silicon carbide,
or contain boron fibers or whiskers.

Generally, aluminum and graphite Epokin primary structural materials are used in areas with lower internal temperatures, such as the armor of aerospace vehicles, or to protect equipment such as avionics. For example, in electrical or hydraulic equipment where equipment inside an aerospace vehicle can withstand higher temperatures, e.g., temperatures on the order of 600'F, the structural material may consist of graphite polyimide. The primary structural materials that can withstand well It can be a metal sheet containing reinforcing fibers such as silicon fibers and in some cases boron uA fibers.The type of fibers used depends on the particular application of the aerospacecraft structural material and the maximum allowable for the vehicle. It depends on the internal temperature. Thus, for example, aluminum lithium alloys and graphite epoxy primary structural materials can withstand temperatures on the order of about 300° F. Titanium aluminide alloys have a temperature resistance capability of 1200° to 1500° F.

If the main structure is subjected to tensile loads, it is advantageous to have fiber reinforcement, such as in composite or metal matrix materials.

The thermal protection system realizing the ceramic matrix outer face sheet 10 and the rigid insulator core 12 of foamed ceramic can be applied to the face by suitable means depending specifically on the composition of the main structural materials, such as bonding, brazing or chemical vapor deposition. It is connected or integrally coupled to the main structure 15 via the sheet 16 . Thus, using a graphite-epoxy composite or aluminum or its alloys as the structural material, the epoxy or polyimide is used to bond the insulator core 12 of, for example, silicon carbide foam ceramic to the face sheet 16 of the composite structural material 15. . Using titanium aluminide, or steel or Inconel matrix materials as the primary tMffi material,
Such materials are suitably brazed to a ceramic insulator core 12, such as foamed silicon carbide, to obtain higher bonding temperatures, since organic bonding of such high temperature materials is feasible. This is because it is not.

It is noted that facesheets 10 and 16 are used on either side of ceramic core 16. thus,
A thermal protection system or unit formed, for example, from an integral silicon carbide face sheet 10 and a silicon carbide foam core 12 is joined to a basic structural fuselage, indicated at 15, which forms, for example, the outer skin of the vehicle. thus,
The outer skin of such carrier 16 becomes the inner facesheet of the clamp consisting of members 10.12 and 16.

The structural component material 14 is attached to the members 10.12 and 10.12 via the inner facesheet 16 by, for example, bonding, brazing, or mechanically as described above, depending on the composition of the particular primary structural component material 14. 16 can be attached to a thermal protection system formed from 16.

Thus, it is the bonding or brazing operation described above that provides the inner facesheet 16 attached to the insulator core 12. This allows the rigid foam core 12 with the outer facesheet 10 thereon to support a thin, eg, metal, facesheet or panel 16 without buckling or wrinkling. Such structural support reduces the thickness of face sheet 16. This structural combination provides the monolithic structure and thermal protection system of this invention.

Figure 2 of the drawings shows a portion of an aerospacecraft forward fuselage structure including an integral structure and thermal protection system (Is/TPS) according to the present invention, which includes a silicon carbide center composite outer skin or face sheet 20. 2, which is bonded to a silicon carbide foam insulation core 22 and to a structural member or blade 24 via a titanium or titanium aluminide inner facesheet 25.

FIG. 3 shows an alternative arc-shaped fuselage structure to that shown in FIG. 2, but the primary structure is a titanium aluminide truss core member 28 having an essentially sinusoidal shape, which is comprised of a titanium aluminide face sheet. 30 is diffusion bonded and it is high temperature brazed by the material silicon carbide foam core 2
2. In this way, the main truss core structural member 2
8 is integrally connected to the foam core 22 and outer facesheet 20 via facesheet 30 to form a unitary structure and thermal protection system.

The concepts of the invention may also be applied to other structural airframe components of an aerospace vehicle, including wings and vertical tails. Thus, for example, graphite epoxy or graphite polyimide wing and tail structural members include a silicon carbide outer facesheet with a foamed silicon carbide core that is integrally bonded to various graphite epoxy structural members by the epoxy or polyimide. It can be thermally stabilized by the thermal protection system of this invention.

In FIG. 4, for example, there is shown a cross-section of a wing 32 having an internal structure that generally includes graphite polyimide girders 34 and ribs 36, which includes a thermal barrier formed by an outer silicon carbide facesheet 38 and a sandwich that includes a foamed silicon carbide core 42. A protection system 40 covers its exterior surface, and the core 42 is suitably joined to the beams and ribs by a graphite polyimide inner facesheet 43 to form a unitary structure and thermal protection system.

In FIG. 5, a vertical stabilizer structure 44 is shown formed of graphite polyimide girders 46 and ribs 48 with a thermal protection system in the form of a set of outer sandwich panels 49 shown in FIGS. 6A and 6B. Each panel 49 is comprised of silicon carbide outer and inner facesheets 50 and 52 and a foamed silicon carbide core 54. In this example, the inner Ω1
1 face sheet 52 is suitably integrally bonded to the struts and ribs of the internal structure of tailplane 44 by polyimide in accordance with the present invention.

Various modifications can be made to this invention. Thus, for example, outer facesheet 10 may be formed of a ceramic matrix material other than silicon carbide, such as silicon nitride.

From the foregoing, it can be seen that the present invention is based on transatmos
It has been found to provide a novel, simple, durable lightweight thermal protection system for each major airframe component of a high-speed aerospace vehicle, such as an ansata+osphcrie) vehicle, which includes an integral part of such components.

Various further modifications of this invention will occur to those skilled in the art, and the invention is not to be construed as limited except as in the appended claims.

[Brief explanation of the drawing]

FIG. 1 shows a cross-sectional view of a monolithic structure and thermal protection system according to the invention. FIG. 2 is a cut away perspective sectional view of a thermal protection system according to the present invention, integrally formed with a component of an aerospace fuselage structure. FIG. 3 is a cut away perspective cross-sectional view similar to FIG. 2, showing the thermal protection system of the present invention integrally connected to different structural components of the fuselage; FIG. 4 is a cross-sectional view of an aerospacecraft wing embodying the thermal protection system of the present invention. FIG. 5 is an elevational perspective view of an aerospacecraft vertical tail structure incorporating the thermal protection system of the present invention. 6A is an elevational perspective view of one of the outer thermal insulation panels of the tail structure of FIG. 5; FIG. FIG. 6B is an enlarged partial view of the thermal protection panel of FIG. 6A taken along circular arrow 6a-6a of FIG. 6A. In the figure, 10 is an outer face sheet, 12 is an insulator core, 14 is another component, 15 is a structural member, 16 is an inner face sheet, 20 is an outer face sheet, 22 is an insulator core, 24 is a structural member or Blade, 25 inner face sheet, 28 truss core member, 30 face sheet, 32 wing, 34 bald, 36 rib, 38
is the outer face sheet, 40 is the thermal protection system, 42 is the core, 43 is the inner face sheet, 44 is the vertical tail structure, 46 is bald, 48 is the rib, 49 is the outer clamping panel, 5
0 is the outer face sheet, 52 is the inner face sheet,
54 is a core.

Claims (1)

Claims: (1) A thermal protection system for a high-speed aircraft, comprising: an outer sheet formed of a ceramic matrix; and a rigid insulator core consisting essentially of a porous foamed ceramic; A thermal protection system, wherein the seat is integrally connected to the core and further includes an internal support structure of the aircraft, the core being suitably coupled to the internal support structure. 2. The thermal protection system of claim 1, wherein the outer sheet comprises silicon carbide. (3) The thickness of the outer sheet is about 0.01" to 0.0
The thermal protection system of claim 1, wherein the internal support structure is connected to the core by an internal sheet forming part of the internal structure, and the internal sheet is joined to the core. The thermal protection system of claim 1, wherein: (5) the insulator core comprises a member selected from the group consisting of silicon carbide and silicon nitride. The thermal protection system of claim 1, wherein the core density ranges from about 0.5 to about 10 pounds per cubic foot. (7) The core thickness ranges from about 0.5" to about 6". The thermal protection system of claim 1. (8) The internal support structure comprises a member selected from the group consisting of composite materials, metals and/or metal matrices including fiber reinforcement. Thermal protection system. (9) the composite material is selected from the group consisting of epoxy and polyimide reinforced with graphite or boron fibers, the metal is selected from the group consisting of aluminum and its alloys and titanium and its alloys; 9. The metal matrix is selected from the group consisting of beryllium, aluminum and titanium and their alloys, steel, molybdenum alloys, and nickel alloys, and the metal matrix comprises graphite, silicon carbide or boron fibers.
Thermal protection system described in. 10. The thermal protection system of claim 1, wherein the core is integrally coupled to the internal support structure by bonding, brazing, or chemical growth. (11) The internal support structure is part of a fuselage, wing, or tail structure of an aircraft, and the core is attached to the internal support structure by an inner facesheet forming an outer skin of the internal support structure. 10. Thermal protection system according to 10.